1. Field of Invention
This invention relates to method and apparatus for spacecraft attitude control, and particularly to satellite transient attitude steady state attitude control during transfer orbit, stationkeeping, synchronous orbit and in general during any change in orbit velocity maneuver. In particular, the invention relates to thruster prebias control in order to compensate for imbalance and cross couplings due to misalignment among a plurality of attitude control thrusters of a spacecraft, such as a synchronous satellite during orbit adjustment.
Spacecraft, particularly satellites designed for geosynchronous orbit and having large structural arrays of solar panels, are particularly susceptible to structural oscillation due to the transient disturbances induced by small rocket engines known as thrusters. The transient disturbances cause not only structural oscillations but can cause undesirable changes in spacecraft attitude which affect antenna direction and, therefore, signal target and signal strength. The problem of spacecraft dynamics and disturbance torque is particularly acute during the orbit adjustment in the stationkeeping mode wherein the satellite attitude is controlled exclusively by gas thrusters.
2. Description Of The Prior Art
The closest known technique for the stabilization of a spacecraft to compensate for transients is described in a research study by Aerospatiale Cannes of France, entitled "COMPATIBILITE' DE LA STABILISATION H. SAT AVEC UNE MISSION TV DIRECTE", Document No. 272/882 A, dated Dec. 25, 1978 by P. Brunet. In this study, a technique was proposed wherein the attitude sensor of the satellite would be offset biased to compensate for the transients induced during the application of thrust. This technique presupposes that there will be attitude error induced which can be calibrated to thrust. However, such a technique is difficult to calibrate because of controller compensation delays. An inherently fast, i.e., wide bandwidth feedback loop would be required to render such a scheme functional. This has a number of inherent limitations and shortcomings such as susceptibility to nonlinearities of the thruster controller, substantial problems due to the uncertainties in the spacecraft dynamic parameters and undesired sensitivity to time delay and sensor phase lag.
The following further patents were uncovered in a search of the prior art:
U.S. Pat. No. 4,325,124 to Renner discloses a system for controlling the direction of the momentum vector of a geosynchronous satellite. The Renner system compensates for disturbance torques applied to a satellite in a way which eliminates the requirement for a thruster control loop. The disturbance torque is employed as a compensating torque to superimpose an artificial misalignment on the incidental misalignment of the satellite's solar panels. The artificial misalignment causes the momentum vector of the satellite to be adjusted to the desired direction to restore the desired attitude. The technique, involving correction which uses solar pressure to correct for solar pressure-induced, misalignment, should not be confused with the present invention.
U.S. Pat. No. 4,174,819 to Bruederle et al. describes a controller for attitude stabilization of a satellite in which the controller generates signals to thrusters and includes correction capabilities to permit efficient attitude control. Specifically, the system employs a two thruster pulse dead-beat mechanism to minimize spacecraft nutation.
U.S. Pat. No. 3,572,618 to Willett describes a method for stabilizing aircraft and missiles in which signals representative of the actual state of an airframe are generated by transducers and compared with a command signal fed to a control system whereby the state of the airframe is controlled. The technique involves modifying the sampling rate with respect to the bending mode of frequencies of sampled data systems.
U.S. Pat. No. 3,490,719 to Cantor et al. describes an attitude control system for providing reliable unidirectional transfer between a coarse mode and a fine mode of target acquisition.
U.S. Pat. No. Re. 30,429 to Phillips discloses a technique for minimizing spacecraft nutation due to disturbing torques. A signal responsive control system operates the attitude or orbit-control forces or a combination of both to minimize spacecraft nutation. In particular, the technique of two-pulse dead-beat control is extended to the use of reaction torque from reaction wheels.
U.S. Pat. No. 4,023,752 to Pistiner et al. discloses the elimination of residual spacecraft nutation due to repulsive torques. A signal responsive control system operates repulsive forces of a spacecraft for a predetermined time corresponding to an integral number of nutation periods in order to eliminate spacecraft nutation.
U.S. Pat. No. 3,937,423 to Johansen discloses nutation and roll-error-angle correction. The system employs two-pulse dead-beat thruster control which includes feedback correction paths by which jet triggering thresholds can be adjusted.
U.S. Pat. No. 4,188,666 to Legrand et al. describes a method and system for torque control and energy storage in a spacecraft. The system is operative to control inertia and reaction wheels to achieve a desired physical orientation and spacecraft activity. It is an object of this invention to minimize disturbance torques. In this system, it is stored kinetic energy which is regulated for controlling a momentum wheel and spacecraft attitude.
U.S. Pat. No. 3,968,352 to Andeen describes a torque control system using a closed tachometer loop to minimize wheel drag torque, bearing noise and the like and to derive an improved torque command for a wheel control system. The torque control system employs pulse-width modulation to control reaction wheels. An error signal derived from the difference between an integrated torque-command signal and angular speed signal is coupled to the motor of a reaction wheel to produce a compensated output torque referenced to disturbance torques.
U.S. Pat. No. 3,624,367 to Hamilton et al. discloses a self-optimized and adaptive attitude control system employing on-off reaction thrusters. Pulse-width modulation is determined by measuring angular velocities of the vehicle in order to minimize the number of propulsion system activations and thereby to prolong thruster life. Means are provided for introducing compensation for the presence of a bias force. The control system does not contemplate elimination of offset error but rather is concerned only with minimizing thruster activity.
U.S. Pat. No. 3,409,251 to Lawson et al. discloses a servo system including compensation for undesirable signals such as those produced by the misalignment, offset and long term drift bias of transducers and error signals in an aircraft.
U.S. Pat. No. 3,330,503 to Love et al. discloses a re-entry guidance system for use with a lifting vehicle entering a planet's atmosphere at high velocity. The system is employed to control aerodynamic control surfaces. During re-entry, a linear accelerometer generates an output which is compared to a precomputed nominal acceleration curve from which an error signal is generated which in turn is added to a command signal to insure that an appropriate amount of force is produced by the associated control surface.